Fast hybrid helicopter with long range with longitudinal trim control

ABSTRACT

A hybrid helicopter includes an airframe provided with a fuselage and a lift-producing surface together with stabilizer surfaces and a drive system including:
         a mechanical interconnection system between a rotor of radius (R) with collective pitch and cyclic pitch control of the blades of the rotor and at least one propeller with collective pitch control of the blades of the propeller; and   at least one turbine engine driving the mechanical interconnection system. The hybrid helicopter includes first members for controlling the angle at which the at least one pitch control surface is set as a function of the bending moment exerted on the rotor mast relative to the pitch axis of the hybrid helicopter, and second members for controlling the cyclic pitch of the blades of the rotor in order to control the longitudinal trim of the hybrid helicopter as a function of flight conditions.

The present invention relates to a rotorcraft with long range withlongitudinal trim control and a high forward speed in cruising flightdue to an optimized lift rotor.

BACKGROUND OF THE INVENTION

More particularly, the invention relates to a hybrid helicopter relatingto an advanced concept for a vertical takeoff and landing aircraft(VTOL).

This advanced concept combines, at reasonable cost, the efficiency invertical flight of a conventional helicopter with the high travel speedperformance made possible by installing modern turbine engines.

In order to understand clearly the object of the invention, it isappropriate to recall the main heavier-than-air craft correspond toairplanes and to rotorcraft.

The term “rotorcraft” is used to designate any vehicle in which all orsome of its lift is provided by one or more propellers of substantiallyvertical axis and large diameter, referred to as rotors or as rotarywings.

There are several distinct types in category of rotorcraft.

Firstly there is the helicopter in which at least one main rotor drivenby a suitable power plant provides both lift and propulsion. Ahelicopter is capable of hovering flight, remaining at a fixed point inthree dimensions, and it can take of and land vertically, and it canmove in any direction (forwards, rearwards, sideways, upwards,downwards). The great majority of rotorcraft produced in the world arehelicopters.

Then there is the autogyro (first made by La Cierva) which is arotorcraft in which the rotor does not receive power, but rotates inautorotation under the effect of the speed of the rotorcraft. Propulsionis provided by a turbine engine or by a propeller having an axis that issubstantially horizontal in forward flight and that is driven by aconventional engine. That configuration is not capable of verticalflight unless the rotor is initially set into rotation by an auxiliarydevice enabling the rotor to be caused to rotate faster: an autogyro istherefore not capable of performing hovering flight but only of movingupwards or downwards on flightpaths having very steep slopes. It is, soto speak, an airplane with a wide range of flying speeds that is notliable to stalling, and that can use short runways.

A gyrodyne is a rotorcraft intermediate between the helicopter and theautogyro in which the rotor provides lift only. The rotor is normallydriven by an engine installation during stages of takeoff, hovering orvertical flight, and landing, like a helicopter. A gyrodyne also has anadditional propulsion system that is essentially different from therotor assembly. In forward flight, the rotor continues to provide lift,but only in autorotation mode, i.e. without power being transmitted tosaid rotor. The Fairey Jet Gyrodyne is an embodiment of this concept.

Several other novel formulae have been studied to a greater or lesserextent, and some have given rise to practical embodiments.

In this respect, mention can be made of the compound rotorcraft thattakes off and lands like a helicopter and that cruises like an autogyro:its rotor, driven in autorotation motion by the forward speed of therotorcraft provides some of the lift, the remainder being provided by anauxiliary wing, a tractor propeller of substantially horizontal axisgenerates the force needed to move in translation. As example, mentioncan be made of the experimental Farfadet SO 1310 compound rotorcrafthaving its rotor propelled by reaction in the takeoff/landingconfiguration and rotating under autorotation in the cruisingconfiguration, propulsion then being provided by a propeller. Thevehicle is provided with two separate turbines for actuating the rotorand the propeller.

The convertible rotorcraft constitutes another particular rotorcraftformula. This term covers all rotorcraft that change configuration whilein flight:

takeoff and landing in a helicopter configuration; and cruising flightin an airplane configuration, e.g. with two rotors being tilted throughabout 90° to act as propellers. The tilting rotor concept has beenimplemented in the Bell Boeing V22 Osprey, for example.

Of those various kinds of rotorcraft, the helicopter is the simplest,and as a result it has been successful in spite of having a maximumhorizontal speed in translation of about 300 kilometers per hour (km/h)which is small and less than that which can be envisaged by compound andconvertible type formulae that are technically more complex and moreexpensive.

Under such conditions, improvements to the above formulae have beenproposed for increasing rotorcraft performance, but without that leadingto solutions that are complicated, difficult to manufacture and tooperate, and consequently expensive.

Thus, patent GB-895 590 discloses a vertical takeoff and landingaircraft comprising the following main elements:

-   -   a fuselage and two half-wings, one on either side of the        fuselage),        -   a horizontal stabilizer and rudder control,        -   at least four interconnected drive units,        -   a main rotor,        -   at least two reversible pitch propellers that are variable            relative to each other, and    -   means under pilot control for transmitting drive power        continuously or from time to time to the rotor and to the        propellers.

Under such circumstances, the main rotor is rotated by the power unitsduring takeoff and landing, during vertical flight, and for horizontalflight at low speed. At high speed, the rotor turns freely without powerbeing transmitted thereto, like an, autogyro, the rotor shaft beingfitted with decoupling means.

U.S. Pat. No. 3,385,537 discloses a helicopter comprising inconventional manner a fuselage, a main rotor, and a tail rotor. The mainrotor is rotated by a first power unit. That vehicle is also fitted withtwo other engines, each engine being disposed at the outermost end oftwo half-wings disposed on either side of said fuselage. The patentrelates to automatically varying the pitch of the blades as a functionof the acceleration exerted on the vehicle while maneuvering or duringgusts of wind, for example, so as to maintain a proper distribution oflift between the rotor and the half-wings. As a result, thecorresponding device contributes to increasing the horizontal speed ofthe rotorcraft by reducing the risks of the blades stalling,constituting sources of variation and damage to the mechanicalassemblies and structures.

U.S. Pat. No. 6,669,137 describes an aircraft fitted with a rotary wingfor operating at very low speed. At high speeds, the rotary wing isslowed down and then stopped, with lift then being produced by ascissors wing. At maximum speeds, the rotary wing and the scissors wingare put into a determined configuration so as to form a kind ofswept-back wing.

The rotorcraft according to U.S. Pat. No. 7,137,591 has a rotor rotatedby a power unit, in particular for takeoff, landing, and verticalflight. A thrust propeller is used in cruising flight, with lift beinggenerated by autorotation of the rotor, possibly with assistance from anauxiliary wing. Furthermore, the rotor mast can be tilted forwards andrearwards a little so as to eliminate the effects due to changes in theattitude of the fuselage that might harm the performance of therotorcraft by increasing its aerodynamic drag.

U.S. Pat. No. 6,467,726 discloses a rotorcraft comprising at least:

-   -   a fuselage,    -   two high wings,    -   at least four propulsion propellers,    -   at least two main rotors without cyclic pitch control, each        connected to one of the two wings,    -   two engines, and the associated means for transmitting power to        the rotors and to the propellers, and    -   a collective pitch control system for each propeller and for        each rotor.

In cruising flight, lift is developed by the two wings, so that the liftdue to the rotor is eliminated either by decoupling the rotor via aclutch provided for this purpose, or by appropriately setting thecollective pitch of the rotor blades.

U.S. Pat. No. 6,513,752 relates to a rotorcraft comprising:

-   -   a fuselage and a wing,    -   two variable-pitch propellers,    -   a rotor with “end” masses,    -   a power source driving the two propellers and the rotor,    -   control means for adjusting the pitch of the propellers so that:        -   in forward flight the thrust from the propellers is exerted            towards the front of the rotorcraft, and        -   in hovering flight, the antitorque function is provided by            one propeller providing thrust towards the front and the            other propeller towards the rear of the rotorcraft, with the            rotor being driven by the power source,    -   the power source comprises an engine and a clutch that, by        disconnecting the rotor from the engine, enables the rotor to        turn faster than an outlet from said engine, because of the        above-mentioned masses.

It is also specified that the clutch permits autogyro mode in forwardflight. In addition, a power transmission gearbox disposed between thepower source and the propellers enables said propellers to operate at aplurality of speeds of rotation relative to the speed of an outlet fromsaid power source.

In the prior art, it is appropriate finally to cite patent applicationUS-2006/0269414 A1, which deals with the particular problem of improvingthe performance of a helicopter both during vertical flight and duringcruising flight. A high speed of rotation for the rotor is then desiredduring vertical flight in order to increase lift, whereas in cruisingflight, said speed of rotation can be reduced while increasing theforward speed of the helicopter.

Consequently, the invention of patent application US-2006/0269414 A1refers more precisely to a main gearbox associated with a second powergearbox driven by the engine installation. The second gearbox includes aclutch device which, when engaged, entrains the main gearbox at a firstspeed of rotation, with disengagement communicating a second speed ofrotation thereto that is lower than the first speed of rotation.Naturally, the main gearbox drives the rotor(s).

Nevertheless, from the above considerations, it can be seen thattechnical solutions that tend to improve the performance of a rotorcraftare based essentially on the following proposals:

-   -   operating the rotor at two distinct speeds of rotation relating        firstly to vertical flight and secondly to cruising flight, by        means of a drive system with variable speed ratios between the        engine installation, the rotor, the propeller(s), and the        various component elements of the drive system.    -   operating the rotor in autogyro mode during cruising flight: the        rotor rotates without driving power being delivered, and then        provides some or all of the lift of the rotorcraft, but also        leads to drag that nevertheless leads to a loss of power because        of a low lift/drag ratio, while in contrast the rotor of a        helicopter propels the rotorcraft in the direction desired by        the pilot.

In particular, it is observed that the operation of a rotor inautorotation like an autogyro during cruising flight makes it necessaryin principle to disconnect the shaft for driving rotation of the rotorfrom the entire power transmission system.

Consequently, this separation is obtained by means such as a clutchhaving the sole function of preventing the rotor being rotated by thepower source(s), and to do so only during the transition from verticalflight to cruising flight.

A device of that type therefore implies additional weight and cost, andconstitutes a penalty in terms of safety.

-   -   Stopping the rotor and reconfiguring it, e.g. a three-blade        rotor stopped in a particular configuration serves as a        swept-back wing for flight at a high forward speed, or indeed,        after stopping, it is possible to envisage folding the rotor        over the fuselage during a rotorcraft-to-airplane transition        stage.

It can be understood that those solutions complicate the technicalimplementation and contribute to increasing weight, and thus toincreasing installed power and consequently expense, but without thatachieving an optimized vehicle.

OBJECTS AND SUMMARY OF THE INVENTION

An object of the present invention is to propose a hybrid helicopter,also sometimes referred to below as a “vehicle”, with longitudinal trimcontrol that makes it possible to overcome the above-mentionedlimitations.

Preferably, the hybrid helicopter must be capable of performing missionseffectively over long periods of vertical flight and also of performingcruising flight at high speed, while also enabling long ranges to beused.

In this respect, the various examples of performance and numerical datacorrespond to concrete and illustrative applications, but should notunder any circumstances be considered as being limiting.

Under such circumstances, a typical mission corresponds for example totransporting 16 passengers at 220 knots (kt) in a vehicle having aweight of about 8 metric tonnes (t) on takeoff over a distance of 400nautical miles (n.miles) at an altitude of 1500 meters (m) under ISAconditions specifying standard atmosphere.

Such performance is highly exceptional in comparison with theperformance of a conventional helicopter such as the Applicant's AS 332MKII type, even though its performance is already remarkable,specifically, for the same tonnage: a recommended cruising speed of 141kt for a similar range, and a fast cruising speed of 153 kt.

According to the invention, a hybrid helicopter with long range and highforward speed and having the following main elements:

-   -   an airframe, i.e. the general structure of the vehicle,        comprising in particular:        -   a fuselage;        -   a lift-providing surface fastened to the fuselage; and        -   stabilizing and maneuvering surfaces, namely for pitch: a            horizontal stabilizer with at least one pitch control            surface movable relative to the front portion or “horizontal            plane”; and for steering: at least one appropriate            stabilizer; and    -   an integrated drive system constituted by:        -   a mechanical interconnection system between firstly a rotor            with collective pitch and cyclic pitch control of the blades            of said rotor, and secondly at least one solely-propulsive            propeller with collective pitch control of the blades of            said propeller; and        -   at least one turbine engine driving the mechanical            interconnection system;            is remarkable in that it comprises first means for            controlling the angle at which said at least one pitch            control surface is set as a function of the bending moment            exerted on the rotor mast relative to the pitch axis of said            hybrid helicopter, and second means for controlling the            cyclic pitch of the blades of said rotor in order to control            the longitudinal trim of the hybrid helicopter as a function            of flight conditions.

It is the collective pitch and the cyclic pitch of the rotor blades thatare adapted to match the varying speed of rotation of the rotor, as afunction of the flightpath airspeed of the vehicle.

Similarly, the collective pitch of the propellers is controlledautomatically in order to deliver the necessary thrust.

More precisely, during cruising flight, another advantage of theinvention consists in controlling the longitudinal cyclic pitch of therotor so as to maintain the attitude of the fuselage at a pitch angle(or longitudinal trim angle) that is equal to the slope of theflightpath, so as to reduce the angle of incidence of the fuselagerelative to the air to zero, thereby minimizing the drag of saidfuselage. Consequently, and during level cruising flight, thelongitudinal trim of the hybrid helicopter is maintained at a value ofzero. Furthermore, and advantageously, the tilting moment of thefuselage is also adjusted by operating at least one moving pitch controlsurface fitted to the horizontal stabilizer, e.g. by means of anelectric actuator, so as to compensate for any offset in the center ofgravity of said hybrid helicopter: this adjustment is obtained in theorywhen the bending moment exerted on the rotor mast relative to the pitchaxis, and as measured by strain gauges, for example, is reduced to zero.

It is advantageous to be able to adjust or even reduce to zero thetilting moment of the fuselage, since firstly that acts directly on thebending moment in the rotor mast and thus on the fatigue stress itsuffers, and secondly that leads to overall balance of the vehicleresulting from the distribution of power between the propellers and therotor. This distribution has an influence on the overall power balance,since the propellers and the rotors have respective differentefficiencies.

During this operation, the distribution of power between the rotor andthe propellers can vary significantly as a function of the angle ofinclination of the rotor disk because of its contribution to variationsin total drag and in the propulsion of the vehicle. By way of example,the power needed in high speed cruising flight is due mainly to theparasitic drag of the vehicle. At 140 kt, parasitic drag representsapproximately 50% of the total power requirements and can reach 75% at220 kt, i.e. three times the power needed for lift. Efficiency at highspeed thus depends on minimizing parasitic drag, which is why it isadvantageous to control the longitudinal trim of said hybrid helicopter.

It will readily be understood that such control is made possible by thetwo degrees of freedom provided by varying at least one pitch controlsurface and secondly by controlling the cyclic pitch of the rotorblades, these first and second means being independent of each other.

In practice, the pitch control surface can be adjusted manually in asimplified version. It is then necessary to provide the instrument panelwith an indicator of the bending moment exerted on the rotor mast so asto enable the pilot to keep it within a determined range by actingmanually on said movable pitch control surface, or indeed on theelectric actuator.

When the pitch control surface 35 is operated automatically in animproved version, said first means respond by automatically controllingthe angle at which said at least one pitch control surface is set via anelectric actuator for example, thereby adjusting the tilting moment ofsaid hybrid helicopter to a first setpoint value that is preferablyequal to zero.

These first means comprise a computer that controls an electric actuatorturning said at least one pitch control surface through an angle thatadapts the tilting moment of the hybrid helicopter to said firstsetpoint value.

To do this, said computer determines the bending moment exerted on therotor mast relative to the pitch axis as deduced from informationdelivered by sensors, said computer ceasing to move said pitch controlsurface when the bending moment exerted on the rotor mast lies within apredetermined range corresponding substantially to the first setpointvalue, preferably equal to zero, for the pitching moment of said hybridhelicopter.

In other words, the term “first setpoint value for the pitching moment”can designate equally well a specific value or a narrow range of valuesfor said pitching moment, in particular because of the dependency of thebending moment exerted on the rotor mast relative to the pitching momentof the hybrid helicopter.

In addition, since the pitching moment is controlled and preferablyreduced to zero, it is also appropriate to control the longitudinal trimof the hybrid helicopter and in particular to reduce it to zero in orderto minimize parasitic drag. The second means thus adapt saidlongitudinal trim to a second setpoint value, preferably equal to zero.These second means comprise at least a cyclic pitch stick that controlsthe cyclic pitch, in particular the longitudinal cyclic pitch of therotor blades via a swashplate and pitch levers.

Naturally, it should be recalled that the lateral cyclic pitch is alsoinvolved to enable the vehicle to perform yaw maneuvers, with variationsin the collective pitch of the rotor blades serving only to vary thelift of each blade by the same amount.

Naturally, the stabilizer for providing yaw control may advantageouslycomprise in front a non-moving portion or fin, and a rear moving portionor rudder. Clearly, the vehicle can be fitted with a plurality ofstabilizers that are substantially vertical or possibly inclinedrelative to the vertical, each provided with a rudder.

Furthermore and in advantageous manner, the speed of rotation Ω of therotor is equal to a first speed of rotation Ω1 up to a first flightpathairspeed V1 of said hybrid helicopter, and is then reduced progressivelyin a linear relationship as a function of the flightpath airspeed of thehybrid helicopter.

If the speed of rotation of a rotor of radius R of a hybrid helicoptertraveling with a flightpath airspeed V is written Ω, then the resultantspeed of the air at the end of the advancing blade is V+U having a speedU equal to ΩR. Under such conditions, the slope of said linearrelationship is advantageously equal to (−1/R) in a coordinate system inwhich speed V is plotted along the abscissa and speed of rotation Ω upthe ordinate. The Mach number at the tip of the advancing blade is thenkept constant.

In practice, the speed of rotation of the rotor is reduced progressivelydown to a second speed of rotation Ω2 corresponding to a secondflightpath airspeed V2 that is the maximum speed of the hybridhelicopter.

Nevertheless, it will be understood that the hybrid helicopter can flyin cruising flight at a flightpath airspeed that is arbitrary, providingit is less than or equal to the maximum flightpath airspeed, such thatthe speed of rotation Ω of the rotor is equal to its first speed ofrotation below V1, and is then determined by the above linearrelationship between V1 and the second flightpath airspeed or maximumspeed V2.

As explained below, this characteristic is essential in the sense thatit makes it possible to maintain the Mach number of the advancing rotorblades at a value of about 0.85, referred to below as the maximum Machnumber. This value is set so as to remain always below the dragdivergence Mach number at which the drag of the rotor increasesconsiderably, and thereby affects the lift/drag ratio of the vehicle andconsequently its performance, while generating vibration that ispenalizing in terms of comfort, safety, and lifetime of the componentsof said vehicle.

The speed of rotation of the rotor of a rotorcraft is conditioned by thediameter of the rotor because the speed at the tip of a blade is limitedby the person skilled in the art to a speed lying in the range 200meters per second (m/s) to 250 m/s in order to avoid degrading theaerodynamic performance of said rotor.

The airspeed of the tip of the “advancing” blade is equal to theairspeed due to the forward speed V of the rotorcraft plus the airspeedU due to the rotation of the rotor.

Consequently, and at a given speed of rotation of the rotor, anyincrease in the forward speed of the rotorcraft leads to a proportionalincrease in the Mach number, equal to the speed at the blade tip dividedby the speed of sound. As stated above, it is appropriate to maintainthe Mach number less than or equal to the drag divergence Mach numberfor the tip profile corresponding to the appearance of compressibilityeffects in the air flow at the blade tip leading to the above-mentioneddrawbacks.

By way of example and on the basis firstly of a maximum Mach numberequal to 0.85 and secondly a speed of 220 m/s at the blade tip due tothe rotation of a rotor having a diameter of 16 m in hovering flight, itis found that the Mach number of the advancing blade reaches 0.85 whenthe rotorcraft is advancing at a speed equal to 125 kt and at analtitude of 1500 m, under ISA conditions, i.e. 5° C.

It can thus be understood that since the intended maximum speed ofadvance is well above that value, e.g. 220 kt, it is important to remedyany increase in Mach number.

According to the invention, it is thus proposed from 125 kt to reduceprogressively the speed of rotation of the rotor from the first speed ofrotation Ω1 of said rotor to its second speed of rotation Ω2 in order tolimit the airspeed at the tip of the advancing blade, e.g. to 171 m/s ata flightpath airspeed of 220 kt, so as to maintain the Mach number atthe tip of the advancing blade at 0.85.

Naturally, this reduction in the speed of rotation of the rotor isaccompanied by a drop in the lift of said rotor. Consequently, the wingcompensates for this drop in lift so as to contribute 31% of the lift at220 kt, as mentioned above in the context of a particular exampleapplication.

It should also be observed that the wing generates lift regardless ofthe forward speed of the rotorcraft, except when hovering, in which caseit presents a peculiar effect of “negative lift” associated withaerodynamic interaction between the rotor and said wing.

Consequently, the lift from the rotor in cruising flight isadvantageously controlled by a suitable, preferably-automatic device forcontrolling the collective pitch while complying with the speed ofrotation of the rotor using setpoint values ranging to said speed ofrotation of the rotor.

Under such conditions, and for the particular version under study, thespeed of rotation Ω of the rotor is equal to a first speed of rotationΩ1 of about 260 revolutions per minute (rpm) up to the first forwardspeed, more correctly referred to as a first flightpath airspeed V1, ofabout 125 kt. Above that speed and up to a second flightpath airspeed ofabout 220 kt, the speed of rotation of the rotor is reducedprogressively to a second value Ω2 of about 205 rpm.

Preferably, the following values are thus used:

-   -   first speed of rotation Ω1 of the rotor: 263 rpm;    -   second speed of rotation Ω2 of the rotor: 205 rpm;    -   first flightpath airspeed V1: 125 kt; and    -   second flightpath airspeed V2: 220 kt.

This preferred solution corresponds to a maximum lift/drag ratio of therotor of not less than 12.2 in the range 150 kt to 220 kt, with thelift/drag ratio of the rotor and the wing taken together exceeding 12above 150 kt.

Consequently, it is ensured that the Mach number at the tips of theadvancing blades is less than 0.85 up to the first flightpath airspeedand is then maintained constant and equal to 0.85 between the first andsecond flightpath airspeeds.

Furthermore, the outlet speeds of rotation of the turbine engine(s), ofthe propeller(s), of the rotor, and of the mechanical interconnectionsystem are mutually proportional, the proportionality ratio beingconstant regardless of the flight configuration of the hybrid helicopterunder normal operating conditions of the integrated drive system.

It can thus be understood that if the hybrid helicopter has only oneturbine engine, it rotates the rotor and the propeller(s) via themechanical interconnection system. If the hybrid helicopter is fittedwith two or more turbine engines, the rotor and the propeller(s) arethen driven in rotation via the mechanical interconnection system bysaid turbine engines.

In other words, the power transmission system operates without anyvariable rotation ratio between the turbine engine(s), the propeller(s),the rotor, and the mechanical interconnection system.

Consequently, and advantageously, the rotor is always driven in rotationby the turbine engine(s), and always develops lift whatever theconfiguration of the vehicle.

More precisely, the rotor is thus designed to provide all of the lift ofthe hybrid helicopter during takeoff, landing, and vertical flightstages, and then to provide part of the lift during cruising flight,with the wing then contributing in part to supporting said hybridhelicopter.

Furthermore, and as described in greater detail below, it is importantto observe that the hybrid helicopter's ability to reach high forwardspeeds makes it necessary to reduce the speed of the air flow at thetips of the blades of the rotor in order to avoid any risk ofcompressibility phenomena in said air flow. In other words, it isnecessary to reduce the speed of rotation of said rotor withoutincreasing its mean lift coefficient, thus leading to a reduction in thelift provided by the rotor.

Thus, the rotor delivers part of the lift to the hybrid helicopter incruising flight, possibly also with a small contribution to propulsionor traction forces (acting as a helicopter) or without any contributionto drag (acting as an autogyro). These conditions of operation thus leadto less power being delivered for the purpose of enabling the rotor toprovide traction. It should be observed that a small contribution topropulsion forces is made by the rotor disk being tilted towards thefront of the vehicle by a small amount only. This process degrades thelift/drag ratio of the rotor very little so it is consequently moreadvantageous in terms of power balance than an additional demand forthrust delivered by the propeller(s).

To do this in cruising flight, the wing provides the additional liftrequired.

Advantageously, the wing is made up of two half-wings, located on eitherside of the fuselage. These half-wings can constitute a high wing, inwhich case they preferably present a negative dihedral angle.Nevertheless, they could also constitute either a low wing, preferablyhaving a positive dihedral angle, or indeed an intermediate wing of anydihedral angle. The shape of these half-wings in plan view maycorrespond, depending on the variant, to half-wings that arerectangular, tapered, forward-swept, swept-back, . . . . Favorably, thespan of the total wing lies in the range 7 m to 9 m for a vehicle havinga maximum authorized takeoff weight of about 8 t.

In a preferred version, the total span of the wing is substantiallyequal to the radius of the rotor, i.e. substantially equal to 8 m, thechord of the wing being set at 1.50 m, i.e. giving an aspect ratio ofabout 5.30. However, these dimensions do not exclude a wing of adifferent aspect ratio.

In a variant of the invention, the wing is fitted with ailerons.

Preferably, the hybrid helicopter is fitted with two propellers locatedon either side of the fuselage, advantageously at the ends of the twohalf-wings.

To provide the required performance for the vehicle, each propeller hasa diameter possibly, but not necessarily, lying in the range 2.5 m to4.5 m, with the diameter of the propellers being 2.6 m in a particularversion that has been studied, as explained below.

Naturally, since the rotor is always driven mechanically by the turbineengine(s), the rotor produces an “opposing rotor torque” tending to makethe fuselage turn in the opposite direction to the rotor. In general,manufacturers install an antitorque rotor at the rear of the fuselage inorder to compensate for the rotor torque. This antitorque rotor in aconventional helicopter is situated behind the fuselage at a distance ofabout 1.5 times the radius of the main rotor, so as to avoid anymechanical interference between them. Such a rotor generally requiresabout 12% of the power of the main rotor in vertical flight. Inaddition, the thrust from said rotor is also used for steering thehelicopter.

Advantageously, the hybrid helicopter of the invention does not have anantitorque rotor, so as to simplify its mechanical assemblies and so asto reduce the weight and the cost of the vehicle, accordingly.

Under such circumstances, the hybrid helicopter is fitted, for example,with at least two propellers, installed on respective half-wings oneither side of the fuselage, with the antitorque and steering controlfunctions being performed by causing the propellers to exertdifferential thrust.

It can be observed that the propellers can be located substantially inalignment relative to the chord plane of the wings or half-wings, orthey can be offset either above or below the wings or half-wings towhich they are connected by a supporting mast.

In other words, in vertical flight, the propeller on the left of thefuselage exerts thrust towards the rear of the vehicle (or “reversethrust”), while the propeller on the right produces thrust towards thefront (or “forward thrust”), assuming that the rotor is turningcounterclockwise when seen from above.

However, the wing span is advantageously of the same order of magnitudeas the radius of the rotor, i.e. as small as possible because of thehigh lift/drag ratio of the rotor in cruising flight, as explainedbelow. As a result the distance between the two propellers is also ofthe same order of magnitude as the radius of the rotor. Under suchconditions, the thrust from the propellers is necessarily greater thanthat from an antitorque rotor.

In addition, and on the basis of the geometrical data given above by wayof example, the diameter of the propellers must be reduced from 3.0 mfor a conventional helicopter to about 2.6 m for the hybrid helicopterso as to allow sufficient space between said rotor and said propellers,thereby further increasing the power needed for the antitorque function.

In any event, this penalty in terms of power is easily compensated bythe large power margin in vertical flight (see below) and by the savingsin weight and cost that result from omitting the antitorque rotor andthe associated power transmission system as represented by horizontaland sloping power transmission shafts and gearboxes known as the“intermediate” and the “rear” gearboxes.

In a variant, the antitorque function can also be implemented in such amanner that, in the above example, the right propeller develops twicethe thrust while the left propeller does not provide any thrust, itbeing understood that under the action of cyclic pitch the rotor mustthen be tilted towards the rear of the vehicle in order to balance thethrust from the right propeller. Under such circumstances, it can beshown that the power required is greater than that needed when the twopropellers deliver thrust in opposite directions.

Naturally, it will be understood that an intermediate solution couldcorrespond to an antitorque function being performed by combining theabove two concepts (pure differential thrust or double thrust with nothrust).

It can readily be understood that because of the constantproportionality ratio between the speeds of rotation of the variouscomponents making up the integrated drive system, that the turbineengine(s), the propeller(s), the rotor, and the mechanicalinterconnection system likewise operate at respective first speeds ofrotation and at respective second speeds of rotation. In other words,first and second speeds of rotation are defined relating respectively tothe first and second flightpath airspeeds, and they are applicable tothe turbine engine(s), to the propeller(s), and to the mechanicalinterconnection system. It should be observed that these second speedsof rotation correspond to 78% of the first speeds of rotation (nominalspeeds: 100% of the first speeds of rotation) in the applicationdescribed above.

Naturally, the speeds of rotation of the turbine engine(s), of thepropeller(s), and of the mechanical interconnection system are reducedprogressively between their respective first and second speeds ofrotation, to comply with the variation in the speed of rotation of therotor between its first and second speeds of rotation in application ofa relationship that is linear or substantially linear.

In this context, it should be recalled that under no circumstances isany use made of variable ratios in speeds of rotation between the engineinstallation, the rotor, the propeller(s), and the various components ofthe integrated drive system.

These various functions described above are made possible by amechanical interconnection system suitable for transmitting power. Sucha system must be capable of transmitting high levels of torque, inparticular because of the high level of power absorbed and therelatively low speed of rotation of the rotor. This requires largereduction ratios in speed of rotation between the various components ofthe drive system, while maintaining a weight that is as small aspossible and ensuring good endurance and good overall safety.

In practice, the mechanical interconnection system comprises thefollowing main components:

-   -   a first main gearbox situated in the fuselage for driving the        rotor at 263 rpm at the nominal speed of rotation (100% of the        first speed of rotation of the rotor), or the first speed of        rotation of the rotor;    -   two second gearboxes for driving the propellers, one gearbox        driving each propeller at 2000 rpm at the nominal speed of        rotation, or first speed of rotation for each propeller;    -   a first shaft, driven by the first gearbox, for driving the        rotor;    -   two second shafts, each being disposed in a respective        half-wing, substantially at one-fourth of its chord, and        delivering power to the rotor and to the propellers, the speed        of rotation of these shafts also being 3000 rpm at the nominal        speed of rotation or first speed of rotation of the second        shafts; and    -   both second shafts are driven by one or more turbine engines by        one or more associated modules which, depending on the type of        turbine engine, reduce the turbine engine speed from 21,000 rpm        or from 6000 rpm to 3000 rpm for the first speed of rotation of        said shafts.

This architecture remains valid for turbine engines whether mounted onthe fuselage or on the half-wings. If mounted on the half-wings, eachspeed reduction module is incorporated in the second gearboxes for thecorresponding propeller instead of being disposed on either side of thefirst main gearbox.

In a basic version, the first main gearbox has two stages, namely:

-   -   a spiral bevel toothed ring driven by two bevel gears, each        connected to one of said second shafts; and    -   said ring acting at the first speed of rotation of the        installation to drive at 1000 rpm the sunwheel of an epicyclic        stage so as to rotate the rotor via planet wheels rotating on a        stationary outer ring.

The two associated modules comprise one or two reduction stagesdepending on the outlet speeds of rotation of the turbine engine. Ingeneral, a single stage suffices for a turbine engine outlet speed of6,000 rpm, whereas two stages are necessary for an outlet speed of21,000 rpm.

Furthermore, the two second gearboxes are fitted with respective speedreduction stages, since the first speed of rotation of the propellers(nominal speed of rotation) is about 2,000 rpm.

Naturally, the number of turbine engines is not limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention and its advantages appear in greater detail in the contextof the description below of embodiments given by way of illustration andwith reference to the accompanying figures, in which:

FIG. 1 is a diagrammatic perspective view of an embodiment of a hybridhelicopter of the invention;

FIG. 2 is a diagrammatic view of the drive system;

FIG. 3 is a diagram of the device for adjusting the longitudinal trim ofthe hybrid helicopter;

FIG. 4 is a diagram showing the relationship for variation in the speedof rotation of the rotor as a function of the forward speed of thehybrid helicopter;

Elements present in two or more distinct figures are given the samereference in each of them.

MORE DETAILED DESCRIPTION

In FIG. 1, there can be seen a hybrid helicopter 1 made in accordancewith the invention.

In the usual way, the hybrid helicopter 1 comprises a fuselage 2 with acockpit 7 at the front thereof, a rotor 10 for driving blades 11 inrotation by means firstly of two turbine engines 5 disposed on top ofthe fuselage 2 (not visible in FIG. 1 because of the presence offairing), on either side of the longitudinal plane of symmetry of therotorcraft, and secondly a main first gearbox MGB (not shown in FIG. 1).

Furthermore, the hybrid helicopter 1 is provided with a high wing 3 madeup of two half-wings 8 disposed on top of the fuselage 2, thesehalf-wings 8 being substantially rectangular in plane view andpresenting a negative dihedral angle.

In a variant of the invention, the high wing 3 is fitted with aileronsnot shown in FIG. 1.

The hybrid helicopter 1 is propelled by two propellers 6 driven by thetwo turbine engines 5, one propeller 6 being disposed at each of theouter ends of the wing 3.

Furthermore, in the vicinity of the rear end of the fuselage 2, surfacesare provided for stabilizing and maneuvering purposes, i.e. for pitchcontrol a horizontal stabilizer 30 having two pitch-control surfaces 35that are movable relative to the front portion 34, and for steering twostabilizers 40, each located at a respective end of the horizontalstabilizer 30.

Specifically, the horizontal stabilizer 40 and the vertical stabilizers50 form a U-shape that is turned upside-down towards the fuselage 2.

Advantageously, the stabilizers 40, which are vertical or inclinedrelative to the vertical, can be constituted by respective non-movingfront portions (or fins) 44 and moving rear portions or rudders 45 foryaw control.

From a dimensional point of view, the hybrid helicopter 1 presentlycorresponds to the following characteristics, relating to a rotorcraftof about 8 t maximum weight authorized for takeoff:

-   -   rotor diameter D: about 16 m;    -   propeller diameter d: 2.6 m;    -   wing span L: 8 m; and    -   aspect ratio λ of the wing: 5.3.

In addition, the hybrid helicopter 1 is fitted with an integrated drivesystem 4 that comprises not only the two turbine engines 5, the rotor10, and the two propellers 6, but also a mechanical interconnectionsystem 15 between these elements as shown diagrammatically in FIG. 2,which is a simplified representation in which it should be understoodthat the rotor 10 and the propellers 6 rotate in planes that aresubstantially orthogonal and not parallel.

With this configuration, the hybrid helicopter 1 is remarkable in thatthe speeds of rotation of the turbine engine outlets, of the propellers,of the rotor, and of the mechanical interconnection system are mutuallyproportional, with the proportionality ratio being constant under normalconditions of operation of the integrated drive system, regardless ofthe flight configuration of the hybrid helicopter.

Naturally, special devices lying outside the ambit to the invention areactivated in the event of possible mechanical breakdowns.

With reference to FIG. 2, the mechanical interconnection systemcomprises the following main components:

-   -   a first or main gearbox MGB situated in the fuselage 2 and        driving the rotor 10 at 263 rpm at the nominal speed of rotation        (or the first speed of rotation of the rotor);    -   two second gearboxes PGB, each of the gearboxes PGB driving one        of the propellers 6 at 2,000 rpm at the nominal speed of        rotation;    -   a first shaft A1 rotated by the first gearbox MGB for driving        the rotor 10;    -   two second shafts A2, each disposed in a respective one of the        half-wings 8, substantially at one-fourth of the chord, and        delivering power to the rotor and to the propellers 6, the        speeds of rotation of the shafts also being 3,000 rpm at the        nominal speed of rotation or first speed of rotation of the        second shafts;    -   the two second shafts A2 are driven in rotation by the two        turbine engines 5 via two associated modules M that, depending        on the type of turbine engine, reduce the speeds of the turbine        engines 5 from 21,000 rpm or from 6,000 rpm to 3,000 rpm for the        first speed of rotation of said shafts.

In a basic version, the first or main gearbox MOB has two stages,namely:

-   -   a spiral bevel toothed ring C1 driven by two bevel gears C2,        each connected to one of said second shafts A2; and    -   said ring C1 operating at the first speed of rotation of the        installation to drive the sunwheel P of an epicyclic stage at        1,000 rpm so as to put the rotor into rotation via planet gears        S rotating on a stationary outer ring CE.

The various above dispositions give the hybrid helicopter 1 thefollowing other characteristics:

-   -   a lift/drag ratio for the rotor of about 12.2 for a flightpath        airspeed greater than 150 kt;    -   a lift/drag ratio F for the rotor and the wing together of about        12, for a flightpath airspeed greater than 150 kt;    -   a maximum speed for the vehicle: 220 kt; and    -   rotor lift: 1.05 times the weight of the helicopter in vertical        flight and lying in the range 0.6 to 0.9 times said weight at        the maximum flightpath airspeed, the rotor 10 being driven        continuously by the turbine engine 5 with power absorption in        cruising flight being reduced to about 500 kw.

Preferably, it should be observed that the ratio of lift of the rotor 10to weight of the hybrid helicopter 1 takes on successively the followingintermediate values:

-   -   0.98 at 50 kt;    -   0.96 at 80 kt;    -   0.90 at 125 kt;    -   0.85 at 150 kt; and    -   0.74 at 200 kt.

In addition, the hybrid helicopter 1 is such that the collective pitchand the cyclic pitch of the blades 11 of the rotor 10 are controlled andadapted as a function of flight conditions.

Concerning the propellers 6, only collective pitch is controlled andadapted as a function of flight conditions.

Furthermore, the hybrid helicopter 1 is adjusted to high-speed cruisingflight so that the rotor 10 exerts lift possibly with a smallcontribution to the propulsion forces but without any contribution todrag. Naturally, this requires power to be absorbed by said rotor 10 tobalance the torque generated by the profile drag and the induced drag ofthe blades 11 of the rotor 10, but this power is relatively small, i.e.about 500 kilowatts (kW) as mentioned above, because of the lift/dragratio of the rotor which is about 12.2 above 150 kt.

The small contribution to propulsion forces occurs because of the rotordisk being tilted a little towards the front of the rotorcraft, whichsolution can be more favorable in terms of power balance than additionalthrust from the propellers because the lift/drag ratio of the rotor isrelatively insensitive to small variations in the trim of the hybridhelicopter.

Furthermore, it is advantageous to be able to adjust the pitching momentof the fuselage since firstly it acts directly on the bending moment inthe rotor mast and thus on the fatigue stressing thereof, and secondlygives rise to the overall balance of the rotorcraft as a result of theway power is distributed between the propellers and the rotor. Thisdistribution has an influence on the overall power balance since thepropellers and the rotor have different respective efficiencies.

As a result and as shown in FIG. 3, the preferably automatic maneuveringof at least one movable pitch-control surface 35 fitted to thehorizontal stabilizer 30 under drive from an electric actuator 70 makesit possible to adjust or even eliminate any pitching moment that resultsfrom the center of gravity becoming offset relative to the lift line ofaction of said hybrid helicopter: this adjustment is obtained when thebending moment of the rotor mast 12 relative to the pitch axis, and asmeasured by strain gauges 71, for example, reduces to zero. In general,this adjustment is relatively slow so that such an actuator is oftencalled a trim actuator.

In addition, and independently, the longitudinal cyclic pitch of therotor 10 is controlled and adapted as a function of flight conditions inorder to maintain the attitude of the fuselage. The longitudinal trimangle is maintained at a value that is equal to the slope of theflightpath, so as to reduce the angle of incidence of the fuselage, andthereby minimizing the drag of the fuselage. Consequently, thelongitudinal trim of the hybrid helicopter 1 is maintained at a value ofzero during a level cruising flight.

In practice, the pitch-control surface 35 can be controlled manually ina simplified version. It is then necessary to provide an indicator onthe instrument panel to indicate the bending moment of the rotor mast12, which bending moment the pilot must then keep within a determinedrange by acting manually on said movable pitch-control surface 35 orindeed on the electric actuator 70.

When the pitch-control surface 35 is maneuvered automatically in animproved version, the electric actuator 70 is controlled by a computer60 that determines the bending moment exerted on the rotor mast 12, asdeduced from information delivered by sensors 71, preferably straingauges disposed on said rotor mast 12. In this way, the computer 60ceases to move said at least one pitch-control surface 35 about its axisAX when the bending moment exerted on the rotor mast relative to thepitch axis lies within a predetermined range corresponding substantiallyto the first setpoint value, for the pitching moment of said hybridhelicopter 1, which value is preferably equal to zero.

As a result, the computer 60, the electric actuator 70, and the sensors71 constitute first means for automatically controlling the angle atwhich said at least one pitch-control surface 35 is set as a function ofthe bending moment exerted on the rotor mast 12 relative to the pitchaxis of the hybrid helicopter 1. Naturally, it is possible to use aplurality of control surfaces 35 for this operation.

Independently, second means (13, 14, 16, 17) control the cyclic pitch ofthe blades 11 of the rotor 10 so as to control the longitudinal trim ofthe hybrid helicopter 1 as a function of flight conditions, adapting itto a second setpoint value relating to said longitudinal trim.

Advantageously, said second setpoint value therefore corresponds to alongitudinal trim angle equal to the value of the slope of theflightpath of the rotorcraft, as mentioned above.

Consequently, this second longitudinal trim value is equal to zeroduring level flight of the hybrid helicopter 1.

Said second means comprise a cyclic pitch stick 13 that controlsservo-controls 14 for imparting the cyclic pitch to the blades 11 of therotor 10 via a swashplate 16 and pitch levers 17.

In practice, it turns out the pilot can use an artificial horizon forensuring a zero longitudinal trim in level flight by using the secondmeans (13, 14, 16, 17).

In contrast, an appropriate system needs to be implemented when anarbitrary slope is required on the flightpath.

For this purpose, said second means are associated with a device 80 forautomatically servo-controlling the longitudinal trim of the hybridhelicopter 1, this automatic servo-control device 80 being integrated inan autopilot 81 and specifically comprising a global positioning system(GPS) 82 for determining said flightpath slope and an attitude andheading reference system (AHRS) 83 for defining the trim of the hybridhelicopter 1 in such a manner as to deduce therefrom the angle ofincidence of the fuselage of said hybrid helicopter 1 relative to theflow of air and to make it zero, together with an anemometerinstallation 84 for correcting errors associated with wind.

This adjustment operation consists in properly positioning the fuselage2 and the rotor 10 at an angle of incidence, that is substantially zerorelative to the flow of air, so as to achieve minimum overall drag andmaximum lift/drag ratio. Compared with operating in autogyro mode, thebalance is favorable, the reason being that for small variations of thelongitudinal trim of the rotor, the lift/drag ratio of the rotor hardlyvaries. Consequently, the rotor is “pulled” without any change inoverall efficiency.

From the point of view of flight mechanics, it is recalled that therotor 10 serves to provide all of the lift of the hybrid helicopter 1during stages of takeoff, landing, and vertical flight, and some of thelift during cruising flight, with the wing 3 then contributing toprovide part of the lift for supporting said hybrid helicopter 1.

Naturally, since the rotor 10 is always driven mechanically by theturbine engine 5, this rotor 10 produces “resistive rotor torque” thattends to cause the fuselage 2 to turn in the opposite direction to therotor 10.

The hybrid helicopter 1 of the invention does not have an antitorquerotor in order to simplify its mechanical assemblies, and consequentlyreduce the weight and the cost of the rotorcraft.

Consequently, since the hybrid helicopter 1 has two propellers 6, eachinstalled on a half-wing 8 on either side of the fuselage 2, thesteering control and antitorque functions are provided by making use ofdifferential thrust, is the difference between the thrusts exerted bythe propellers.

In other words, in vertical flight, propeller 6 on the left fuselageexerts thrust towards the rear of the rotorcraft (“rear thrust”) whilethe propeller 6 on the right produces thrust towards the front (“frontthrust”), assuming that the rotor 10 rotates anticlockwise when seenfrom above.

In a variant, the antitorque function may also be performed in such amanner that, on the above example, the right propeller 6 develops doublethrust while the left propeller 6 does not provide any thrust, it beingunderstood that the rotor 10 must then be inclined towards the rear ofthe rotorcraft in order to balance the thrust from the right propeller.Under such circumstances, it can be shown that more power is requiredthan when the two propellers provide thrust in opposite directions.

Consequently, and on the basis of the above-described example and ofFIG. 4, the speed of rotation of the rotor 10 is equal to a first speedof rotation Ω1 of 263 rpm up to a first forward speed V1, more correctlyreferred to as a first flightpath airspeed of 125 kt. At higher speedsand up to a second flightpath airspeed V2 of 220 kt, the speed ofrotation of the rotor is reduced progressively to a second speed ofrotation Ω2 of 205 rpm. The progressive reduction in the speed ofrotation of the rotor 10 between the first and second flightpath speedsvaries in application of a linear relationship of slope (−1/R) where Ris the radius of the rotor, and in a coordinate system where speed V isplotted along the abscissa and speed of rotation Ω of the rotor (10) isplotted up the ordinate.

The person skilled in the art knows that if airspeed increases, the Machnumber at the end of the advancing blade of a rotorcraft rotor reachesthe Mach number known as the “drag-divergence” Mach number.

Then, and at a maximum Mach number that is less than or equal to theso-called drag-divergence Mach number of the end profile of the blade,the speed of rotation of the rotor needs to reduced progressively as afunction of the increase in the forward speed of the rotorcraft so as toavoid exceeding that limit.

If the speed of sound is written c, the Mach number at the end of theadvancing blade is equal to the expression (V+U)/c or indeed (V+ΩR)/c.Imposing a maximum Mach number equal to Mm amounts to causing Q to varyin application of the following linear relationship [(c.Mn−V)/R].

Assuming that the maximum Mach number is equal to 0.85 and a peripheralairspeed at the blade tip of 220 m/s in vertical flight (speed ofrotation of the rotor 263 rpm), the Mach number of the advancing blades11 reaches 0.85 at a flightpath airspeed of 125 kt at an altitude of1500 meters under ISA conditions (outside temperature: 5° C.).

Over the range 125 kt to 220 kt, the speed of rotation Ω is adapted tocomply with the above-specified relationship.

When the flightpath airspeed of the rotorcraft is 220 kt, the airspeedat the tip of the blade due to rotation is equal to 171 m/s (speed ofrotation of the rotor: 205 rpm or 78% of the nominal speed of therotation of the rotor) and the advance parameter μ is equal to 0.66. Atthis value for the advance parameter, the lift of the rotor cannot bemaintained without a large increase in the chord of the blades (60%increase which would lead to a conventional four-blade helicopter havinga chord of one meter), in order to maintain a mean blade liftcoefficient Czm of less than 0.5 and thus avoid separation at theretreating blade. It is clear that such overdimensioning of the bladesat a high forward speed would lead to a significant increase in theweight of the rotorcraft and to penalizing its performance.Consequently, the rotor of the hybrid helicopter 1, having a maximumauthorized takeoff weight of about 8 t, is progressively taken over by awing 3 of small span L that delivers lift of about 31% at 220 kt. Undersuch conditions, when the flightpath airspeed increases, the liftcoefficient Czm of the blades, which in vertical flight is equal to 0.5(wing lift contribution estimated at 4.5%), decreases to reach 0.43 at125 kt because of the increase in lift from the wing 3, and it increasesto reach a value of 0.54 at 220 kt because of the reduction in the speedof rotation of the rotor to 78% of its nominal speed of rotation. Undersuch conditions, the rotor operates with a maximum lift/drag ratio ofabout 12.2.

Finally, the general architecture of the hybrid helicopter 1 of theinvention, associated with:

-   -   a constant proportionality ratio between the speeds of rotation        of the turbine engines 5, of the rotor 10, of the propellers 6,        and of the mechanical interconnection system 15, the drive        system rotating at a first speed up to a first forward speed of        the vehicle, and then decreasing to a second speed of rotation        to a second forward speed equal to the maximum forward speed;    -   a device 70 for controlling and maintaining the longitudinal        trim of the vehicle at a value of zero and without a tail rotor;    -   an integrated drive system 4; and    -   a mechanical interconnection system 15, preferred over a        jet-propelled rotor for better mechanical efficiency and less        noise;    -   all contribute to obtaining high performance.

Thus, the hybrid helicopter 1 is characterized by exceptionalversatility, enabling it to optimize the compromise between speed,range, and weight of the vehicle. By way of example, and with about 2 tof fuel, it is possible to obtain the following performance for a hybridhelicopter 1 weighing about 8 t and transporting 16 passengers:

-   -   hovering duration: 4.2 hours;    -   usable range at 220 kt: 511 n.miles; and    -   usable range at the economic cruising speed of 125 kt: 897        n.miles.

Similarly, and still by way of example, a 400 n.mile mission at 20 ktcan be performed with about 1.6 t of fuel and 20 minutes reserve.

These results demonstrate the large amount of flexibility andadaptability of the hybrid helicopter 1 and its advantages compared witha conventional helicopter. The cruising speed of conventional helicopterhas only a minor influence on its fuel consumption during a mission, soits maximum cruising speed is relatively close to the economic cruisingspeed, such that the only option for increasing the usable range of thevehicle significantly is to reduce the number of passengers so thatadditional fuel can be taken on board.

Naturally, the present invention can be subjected to numerous variationsas to its implementation. In particular, it is important to observe thatthe invention as described relates in particular to a hybrid helicopterweighing about 8 t in all. Nevertheless, the invention is applicable toany rotorcraft of arbitrary weight, for example from a low weight droneto a vehicle of very large tonnage. Although several embodiments aredescribed above, it will be understood that it is not conceivable toidentify exhaustively all possible embodiments. It is naturally possibleto envisage replacing any of the means described by equivalent meanswithout going beyond the ambit of the present invention.

1. A hybrid helicopter having a long range and a high forward speed, thehelicopter comprising: an airframe comprising: a fuselage; alift-producing surface secured to the fuselage; and stabilizing andmaneuvering surfaces comprising: a horizontal stabilizer with at leastone pitch control surface that is movable relative to a front portion ofthe horizontal stabilizer; and at least one suitable stabilizer; and anintegrated drive system comprising: a mechanical interconnection systemconnected between a rotor of a radius with collective pitch and cyclicpitch control of the blades of said rotor, and at least one propellerwith collective pitch control of the blades of said propeller; and atleast one turbine engine driving the mechanical interconnection system;the helicopter comprising a pitch control surface system for controllingthe angle at which said at least one pitch control surface is set as afunction of the bending moment exerted on the rotor mast relative to thepitch axis of said hybrid helicopter, and a cyclic pitch system forcontrolling the cyclic pitch of the blades of said rotor in order tocontrol the longitudinal trim of the hybrid helicopter as a function offlight conditions, wherein said pitch control surface system and saidcyclic pitch system communicate with a flight computer to automaticallycontrol the pitch control surface and the cyclic pitch of the blades. 2.A hybrid helicopter according to claim 1, wherein said pitch controlsurface system automatically controls the angle at which said at leastone pitch control surface is set as a function of the bending momentexerted on the rotor mast relative to the pitch axis of said hybridhelicopter.
 3. A hybrid helicopter according to claim 2, wherein saidpitch control surface system and said flight computer controls anelectric actuator turning said at least one pitch control surface aboutan axis through an angle that adapts the tilting moment of said hybridhelicopter to said first setpoint value.
 4. A hybrid helicopteraccording to claim 2, wherein, by automatically controlling the angle atwhich said at least one pitch control surface is set, said pitch controlsurface system adjusts the tilting moment of said hybrid helicopter to afirst setpoint value.
 5. A hybrid helicopter according to claim 4,wherein said first setpoint value for the tilting moment of said hybridhelicopter is equal to zero.
 6. A hybrid helicopter according to claim4, wherein the flight computer determines the bending moment exerted onthe rotor mast, as deduced from information delivered by sensors, saidflight computer ceasing to move said at least one pitch control surfacewhen the bending moment exerted on the rotor mast lies in apredetermined range corresponding substantially to the first setpointvalue for the tilting moment of said hybrid helicopter.
 7. A hybridhelicopter according to claim 6, wherein the sensors are strain gaugesplaced on the rotor mast.
 8. A hybrid helicopter according to claim 1,wherein the cyclic pitch system adapts the longitudinal pitch of saidhybrid helicopter to a second setpoint value.
 9. A hybrid helicopteraccording to claim 8, wherein said second setpoint value for thelongitudinal trim is equal to the value of the slope of the flight pathof said hybrid helicopter.
 10. A hybrid helicopter according to claim 8,wherein said second setpoint value for the longitudinal trim of saidhybrid helicopter is equal to zero.
 11. A hybrid helicopter according toclaim 1, wherein said pitch control surface system is independent ofsaid cyclic pitch system.
 12. A hybrid helicopter according to claim 1,wherein said cyclic pitch system comprises a cyclic pitch stick thatcontrols servo-controls for imposing the cyclic pitch of the blades ofthe rotor via a swashplate and pitch levers.
 13. A hybrid helicopteraccording to claim 1, wherein said cyclic pitch system is associatedwith a device for automatically servo-controlling the longitudinal pitchof the hybrid helicopter to the slope of the flight path of said hybridhelicopter, said automatic servo-control device being integrated in anautopilot specifically including a GPS for determining said flight pathslope and an attitude and heading reference system for defining the trimof the hybrid helicopter so as to deduce therefrom the angle ofincidence of the fuselage of said hybrid helicopter relative to the airand reduce it to zero, together with an anemometer installation forcorrecting errors associated with disturbances due to wind.
 14. A hybridhelicopter according to claim 1, wherein said pitch control surfacesystem automatically controls the angle at which two pitch controlsurfaces are set as a function of the bending moment of the rotor mastrelative to the tilting axis of said hybrid helicopter.